Rotor for a turbomachine

ABSTRACT

The invention relates to a rotor ( 10 ) for a turbomachine, in particular for a gas turbine, having a basic rotor body ( 12 ) and a plurality of blades ( 14 ), wherein at least one damping element ( 24 ) is provided in the circumferential direction between adjacent blades ( 14 ) of the rotor ( 10 ).

The invention relates to a rotor for a turbomachine, in particular for a gas turbine, having a basic rotor body and a plurality of blades.

Built-in or integrally bladed rotors for turbomachines are known from the prior art. The denotation of rotors having integral blading depends on whether a rotor or rotor support (called a basic rotor body in the following) is present that is shaped like a disk in cross section (blisk) or is ring-shaped in cross section (bling). Blisk is the abbreviated form of bladed disk and bling of bladed ring.

In particular, integrally bladed rotors are known, during the production of which, the blades will be joined together into a blade ring in a first step and subsequently will be fastened to a basic rotor body. Vibrations can be transferred from one blade to the adjacent blades via the structure of such a blade ring. In addition, a replacing of individual blades in order to repair the rotor is not provided and is difficult for such rotors.

The problem of the invention is to create a rotor for a turbomachine with fewer vibrations.

The problem is solved by a rotor according to the invention for a turbomachine, in particular for a gas turbine, having a basic rotor body and a plurality of blades, wherein at least one damping element is provided in the circumferential direction between adjacent blades of the rotor. In this way, vibrations of a blade can be damped and in particular, the transfer of vibrations to adjacent blades can be reduced.

Preferably, the rotor is an integrally bladed rotor, in particular a rotor in which the basic rotor body and blades are welded directly to one another or via a separate intermediate piece; in particular, they are friction-welded. The use of integrally bladed rotors in turbomachines makes possible a savings in weight when compared with multi-part rotors and is thus particularly of advantage for aircraft engines.

According to a preferred embodiment, the blades have a blade neck, by means of which they are joined to the basic rotor body. The neck region of the blade can have different structural and functional features that can be integrated into the blade as one piece in this way.

The blade necks have, e.g., extensions that together form a shroud that bounds by its radially outward-pointing surface an annular flow channel of the turbomachine. Such a shroud can separate the flow channel from the basic rotor body, so that the basic rotor body is protected against loads, particularly by hot gases. It is possible that a hollow space between adjacent rotors in the turbomachine is formed by the shroud, and this space can be used, in particular, for cooling the rotor.

Preferably, the at least one damping element is essentially disposed in the region of the blade neck. In this way, the damping element does not occupy any structural space in the flow channel of the turbomachine.

According to a preferred embodiment, the blade necks of adjacent blades are distanced from one another in the circumferential direction, and a free space will be formed between the blade necks. Due to this free space, the free structural space around a blade is enlarged, whereby it is possible, for example, to fasten an individual blade to the basic rotor body or to separate this blade from it, independently from the adjacent blades.

The damping element preferably forms a seal of the free space between adjacent blade necks relative to a flow channel of the turbomachine. This prevents the free space between the blade necks from acting negatively on the flow ratios or aerodynamic conditions in the turbomachine.

The damping element can form a part of the shroud and can limit the annular flow channel by its surface that is directed radially outward. The damping element thus makes possible a smoother transition between the extensions of adjacent blades.

One embodiment provides that the damping element is disposed as an inserted piece between the blades. Thus, one inserted piece can be inserted or removed independently from other components of the rotor.

The damping element can be joined to the adjacent blades in a form-fitting manner, via a press connection and/or cohesively. This reinforces the mechanical stability of the rotor.

It is also possible that the damping element is joined to the basic rotor body in a form-fitting manner, via a press connection and/or cohesively.

Preferably, a channel is provided between the at least one damping element and the basic rotor body. Such a channel, for example, makes possible a connection between the front and back sides of the rotor.

The channel can be formed at least partially by a notching in the basic rotor body and/or in the blade neck.

Preferably, the channel is a cooling channel. A cooling of the basic rotor body, the blades, and/or the damping element is made possible in this way.

In particular, a plurality of damping elements that are joined together are provided. For example, this makes possible the coupling of the plurality of damping elements and reduces the number of individual parts in the assembly of the rotor.

According to another embodiment, the plurality of damping elements is introduced on a support ring that can be axially attached. This makes possible a simple assembly by inserting the support ring in the rotor in the axial direction. The support ring can be designed self-supporting, so that its mass does not contribute to the load on the basic rotor body when it is rotating.

Additional features and advantages of the invention result from the following description and from the drawings, to which reference is made. In the drawings:

FIG. 1 shows a section of a rotor according to the invention with blade necks distanced from one another prior to the insertion of the damping elements;

FIG. 2 shows the rotor from FIG. 1 with inserted damping elements;

FIG. 3 shows another view of the rotor according to FIG. 2; and

FIG. 4 shows a detail view of the region of the blade neck of a rotor according to the invention.

FIG. 1 shows a section of a rotor 10 having a basic rotor body 12, on which are fastened several blades 14. The blades 14 have a blade surface 16 and a blade neck 18. On the side lying in the circumferential direction, the blade neck 18 has a recess 20, which makes possible a lightweight construction of the rotor 10 by reducing the weight of the blades 14. In the axial direction, the blade necks 18 have extensions 22 that together form a shroud that bounds by its radially outward-pointing surface an annular gas flow channel of the turbomachine.

Rotor 10 is designed for a gas turbine, whereby it can be disposed in the turbine section or in the compressor section. The invention, of course, can also be applied to rotors of other turbomachines.

The blades 14 are joined to the basic rotor body 12 via the blade necks 18. In the embodiment shown, basic rotor body 12 and blades 14 are welded together directly.

Care must be taken that blades 14 are manufactured particularly from materials that do not permit fusion welding methods, for example, monocrystalline materials. Blades 14 are thus joined to the basic rotor body 12 via friction-welding processes, in particular linear friction-welding processes, or inductive high-frequency press welding.

Blades 14 are distanced from one another in the circumferential direction, particularly in the region of the blade neck 18, whereby a free space is formed between the blades 14. The free spaces on both sides of a blade 14 provide sufficient structural space to fasten an individual blade 14 to the basic rotor body 12 or to separate this blade from it, independently from the adjacent blades 14. This makes possible a simplified production process and particularly a simplified repair process by being able to replace an individual blade 14.

It is also possible that blades 14 are welded to the basic rotor body 12 via a separate (in particular, polycrystalline) intermediate piece. In this way, one need only pay attention to the particular joining method when the intermediate piece and blades 14 are joined.

It is also possible that blades 14 are not integrally joined to the basic rotor body 12, but are anchored in the basic rotor body 12 via blade feet.

FIG. 2 and FIG. 3 show the section of rotor 10 from FIG. 1, whereby damping elements 24 are inserted between the blade necks 18. FIG. 4 shows a detail view of the region of the blade neck with inserted damping elements 24. The damping elements 24 reduce the transfer of vibrations from one blade 14 to adjacent blades 14 and also may have a sealing function.

The damping elements 24 are joined to the adjacent blade necks 18 in a form-fitting manner and/or cohesively and/or via a press fitting. Of course, the damping elements 24 can also be joined correspondingly to the basic rotor body 12.

The damping elements 24 fill the free space between the adjacent blade necks 18 and seal this free space relative to the flow channel of the turbomachine. The shape of the damping elements 24 is adapted to the shape of the blade necks 18, particularly in the region of extensions 22. The radially outwardly pointing surface of the damping element 24 thus bounds the annular flow channel of the turbomachine.

In the embodiment shown, the damping elements 24 are disposed as individual inserted pieces, each between two adjacent blades 14. A plurality of damping elements 24 can also be joined together. In this way, on the one hand, the damping elements 24 can be coupled together, and, on the other hand, the number of individual parts in the production of rotor 10 is reduced. It is also possible that a damping element 24 is not introduced between each adjacent pair of blades.

A plurality of damping elements 24 can be introduced on a support ring that can be fastened axially and that is inserted into the rotor 10 in the axial direction during assembly. Such a support ring is designed self-supporting, so that its mass does not contribute to the load of the rotor 10 or of the basic rotor body 12 during operation.

Channels 26, which are formed by a notches 28 in the basic rotor body 12 and in the blade neck 18, are provided between the damping elements 24 and the basic rotor body 12.

The channels 26 serve as cooling channels because a cooling fluid flows through them. Channels 26 thus make possible a flow of cooling fluid in the axial direction from the axially front side of the rotor 10 to the axially back side of rotor 10.

Ni-based alloys are particularly considered as the material for the damping elements 24. The material should be softer than the adjacent blade necks. 

1. A rotor (10) for a turbomachine, in particular for a gas turbine, having a basic rotor body (12) and a plurality of blades (14), is hereby characterized in that at least one damping element (24) is provided in the circumferential direction between adjacent blades (14) of the rotor (10).
 2. The rotor according to claim 1, further characterized in that the rotor (10) is an integrally bladed rotor (10), in particular a rotor (10) in which the basic rotor body (12) and blades (14) are welded, particularly friction-welded, directly to one another or via a separate intermediate piece.
 3. The rotor according to claim 1, further characterized in that the blades (14) have a blade neck (18), by means of which they are joined to the basic rotor body (12).
 4. The rotor according to claim 3, further characterized in that the blade necks (18) have extensions (22) that together form a shroud that bounds by its radially outward-pointing surface an annular flow channel of the turbomachine.
 5. The rotor according to claim 3, further characterized in that the damping element (24) is essentially disposed in the region of adjacent blade necks (18).
 6. The rotor according to one of claim 3, further characterized in that the blade necks (18) of adjacent blades (14) are distanced from one another in the circumferential direction and a free space is formed between the blade necks (18).
 7. The rotor according to claim 6, further characterized in that the damping element (24) forms a seal of the free space between adjacent blade necks (18) relative to a gas flow channel of the turbomachine.
 8. The rotor according to claim 7, further characterized in that the damping element (24) forms a part of the shroud and bounds the annular gas flow channel by its surface pointing radially outward.
 9. The rotor according to claim 1, further characterized in that the damping element (24) is disposed as an inserted piece between the blades (14).
 10. The rotor according to claim 1, further characterized in that the damping element (24) is joined to the adjacent blades (14) in a form-fitting manner, via a press connection and/or cohesively.
 11. The rotor according to claim 1, further characterized in that the damping element (24) is joined to the basic rotor body (12) in a form-fitting manner, via a press connection and/or cohesively.
 12. The rotor according to claim 1, further characterized in that a channel (26) is provided between the at least one damping element (24) and the basic rotor body (12).
 13. The rotor according to claim 12, further characterized in that the channel (26) is formed at least partially by a notch (28) in the basic rotor body (12) and/or in the blade neck (18).
 14. The rotor according to claim 12, further characterized in that the channel (26) is a cooling channel.
 15. The rotor according to claim 1, further characterized in that a plurality of damping elements (24) that are joined together are provided.
 16. The rotor according to claim 1, further characterized in that a plurality of damping elements (24) that are introduced on an axially attachable support ring is provided. 